The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade and vane cooling.
In a gas turbine engine, air is compressed in a compressor, mixed with fuel and ignited in a combustor for generating hot combustion gases which flow downstream through one or more stages of turbine nozzles and blades. The nozzles include stationary vanes followed in turn by a corresponding row of turbine rotor blades attached to the perimeter of a rotating disk. The vanes and blades have correspondingly configured airfoils which are hollow and include various cooling circuits and features which receive a portion of air bled from the compressor for providing cooling thereof against the heat from the combustion gases which flow therearound.
Turbine vane and blade cooling art is crowded with various features configured for enhancing cooling and reducing the required amount of cooling air for increasing the overall efficiency of the engine while obtaining a suitable useful life for the vanes and blades. For example, typical vane and blade airfoils in the high pressure turbine section of the engine include variously configured cooling holes which extend through the pressure side, or suction side, or both, for discharging a film of cooling air along the outer surface of the airfoil to effect film cooling in a conventional manner.
Since the film cooling air is being discharged from inside the airfoils to outside the airfoils over which the combustion gases flow, a suitable differential pressure must be provided for preventing backflow of the combustion gases into the airfoils. However, excessive differential pressure of the cooling air relative to the combustion gases decreases the effectiveness of the film cooling holes as evaluated by a conventional blowing ratio of the density and velocity of the cooling air relative to the density and velocity of the combustion gases of sufficient strength to blow the coolant film off the airfoil surface downstream of the holes.
It is desirable to reduce the blowing ratio to a suitable value for maximizing performance of the film cooling air while providing sufficient backflow margin to prevent ingestion of the combustion gases into the airfoils during operation.
A common film cooling hole is in the form of a cylindrical aperture inclined axially through the airfoils sides, such as the pressure side, for discharging the film air in the aft direction. The film cooling holes are typically provided in a radial or spanwise row of holes at a specific pitch spacing therebetween. In this way, a row of the film cooling holes discharges corresponding cooling films which form an air blanket for protecting the outer surface of the airfoil from the hot combustion gases during operation.
In the region of the blade leading edge, it is also known to incline the cylindrical film cooling holes at an acute span angle to position the hole outlets radially above the hole inlets and discharge the cooling film radially outwardly from the respective holes.
In order to improve the performance of film cooling holes, it is also conventional to modify their shape to effect diffusion which reduces the discharge velocity of the airflow therethrough and increases static pressure thereof. Diffusion film cooling holes are found in many patented configurations for improving film cooling effectiveness with suitable blowing ratios and backflow margin. A typical diffusion film cooling hole may be conical from inlet to outlet with a suitable increasing area ratio therebetween for effecting diffusion without undesirable flow separation. In this way, diffusion occurs in three axes, i.e. along the length of the hole and the two in-plane orthogonal axes perpendicular thereto.
Other types of diffusion film cooling holes are also found in the prior art including various rectangular shaped holes having different performance. Like the conical diffusion holes, the rectangular diffusion holes also effect diffusion in three dimensions as the cooling air flows therethrough and is discharged along the outer surface of the airfoil.
As indicated above, the various diffusion film cooling holes are typically arranged in rows extending along the span or radial axis of the airfoil, and are positioned closely together as space permits for collectively discharging film cooling air. Since a suitable space must be provided between the adjacent film cooling holes for maintaining suitable strength, for example, the discharge film cooling air does not provide 100% coverage along the span line of the corresponding row of holes.
For example, a typical hole pitch spacing is ten diameters of the circular hole inlet. In the example of the spanwise inclined cylindrical film cooling holes described above, a typical span angle is about 30.degree., with a 0.25 mm diameter. And, the effective coverage of the row of film cooling holes may be defined by a coverage parameter represented by the span height of the cooling hole along the airfoil outer surface divided by the pitch spacing of adjacent holes. For an inclined cylindrical hole, the outer surface span height of the hole is simply the diameter of the hole divided by the sine of the inclination angle. This results in a 20% coverage value for 30.degree. inclined cylindrical holes at a ten diameter spacing.
This coverage may be compared with a row of conical diffusion holes having 0.25 mm circular inlets increasing in area to circular outlets having a diameter of about 0.46 mm, with the same centerline spanwise hole spacing or pitch of ten inlet diameters. The corresponding coverage value is 36%, which is an improvement over the simple cylindrical holes.
Accordingly, it is desired to further improve film cooling coverage in a row of diffusion holes within the available space while maintaining blade strength.